The manufacture of a ceramic matrix composite (CMC) part typically includes laying up pre-impregnated composite fibers having a matrix material already present (prepreg) to form the geometry of the part (pre-form), autoclaving and burning out the pre-form, infiltrating the burned-out pre-form with the melting matrix material, and any machining or further treatments of the pre-form. Infiltrating the pre-form may include depositing the ceramic matrix out of a gas mixture, pyrolyzing a pre-ceramic polymer, chemically reacting elements, sintering, generally in the temperature range of 925 to 1650° C. (1700 to 3000° F.), or electrophoretically depositing a ceramic powder. With respect to turbine airfoils, the CMC may be located over a metal spar to form an outer covering over the metal spar or to form only the outer surface of the airfoil.
Examples of CMC materials include, but are not limited to, carbon-fiber-reinforced carbon (C/C), carbon-fiber-reinforced silicon carbide (C/SiC), silicon-carbide-fiber-reinforced silicon carbide (SiC/SiC), alumina-fiber-reinforced alumina (Al2O3/Al2O3), or combinations thereof. The CMC may have increased elongation, fracture toughness, thermal shock, dynamic load capability, and anisotropic properties as compared to a monolithic ceramic structure.
CMC hollow airfoils may suffer from peel stresses in the trailing edge (TE) and/or inter-laminar tensile (ILT) wishbone stresses at the TE-core interface. Due to the pressurization of the cavity within the CMC hollow airfoil during service, the leading edge (LE) and TE are subjected to high ILT ballooning stresses. Since the LE and TE cores are rounded on the interior at the cavity, a radial roll of matrix and fiber is placed to take up space where it is not possible to contain normal CMC plies.